Turbomachine aerofoil and a method of production

ABSTRACT

A cast and hot isostatically pressed gamma titanium aluminide turbine blade (26) has an aerofoil portion (38) including a concave pressure surface (44) and a convex suction surface (46). The concave pressure surface (44) has surface irregularities (48), produced by the action of the hot isostatic pressing process on voids (56) located within the cast turbine blade (26), located in that region (50) of the concave pressure surface (44) where in operation aerodynamic separation (G) occurs to minimise the aerodynamic effects of the surface irregularities (48) upon the operation of the turbine blade (26). The remaining portion of the concave pressure surface (44) and the whole of the convex suction surface (46) are substantially free of surface irregularities (48).

This is a division of application Ser. No. 08/495,174, filed Jun. 27,1995, now U.S. Pat. No. 5,609,470.

FIELD OF THE INVENTION

The present invention relates to turbomachine aerofoils, and isparticularly concerned with turbomachine aerofoils which have beenproduced by casting and hot isostatic pressing processes.

BACKGROUND OF THE INVENTION

It is well known in the art to produce turbomachine aerofoils, whether aturbomachine blade or a turbomachine vane, by casting molten metal intoa mould and allowing the molten metal to cool in the mould and to takethe shape defined by the mould. One of the problems associated with thecasting process is that there is a possibility of voidage, or porosity,in the finished cast component. It is also well known in the art toremove this unwanted voidage, or porosity, by the use of a hot isostaticpressing process. This involves placing the finished cast component intoan autoclave, i.e. a high pressure vessel, sealing the autoclave andapplying high pressures and high temperatures on the cast components inthe autoclave.

A problem with metals, or alloys, which exhibit a high degree ofvoidage, or porosity, for example gamma titanium aluminides is thatsurface irregularities, ie, dimples may result on the surface of thecast and hot isostatically pressed component. In the case ofturbomachine aerofoils, the surface irregularities may have detrimentaleffects on the performance of the turbomachine aerofoils in operation.

One way to overcome the problem is to cast the component oversize andthen to machine the cast and hot isostatically pressed turbomachineaerofoil to exact size and shape. However, this machining operation isvery expensive. Another way to overcome the problem is to cast thecomponent to size and shape and to fill in the surface irregularities onthe cast and hot isostatically pressed turbomachine aerofoil by forexample weld deposition.

SUMMARY OF THE INVENTION

Accordingly the present invention seeks to provide a cast and hotisostatically pressed turbomachine aerofoil having surfaceirregularities in which the effects of the surface irregularities arereduced or minimised.

Accordingly the present invention provides a cast and hot isostaticallypressed turbomachine aerofoil including a concave surface and a convexsurface, the concave surface having surface irregularities produced bythe action of the hot isostatic pressing process on voids in the castturbomachine aerofoil located in that region of the concave surfacewhere in operation aerodynamic separation, a slow flow or a stagnantflow occurs, the remainder of the concave pressure surface and theconvex surface being substantially free of surface irregularities.

Preferably the turbomachine aerofoil includes a shroud located at oneend of the aerofoil, the shroud has a recess in its surface facing awayfrom the aerofoil.

Preferably the turbomachine aerofoil is a blade, or a vane, for a gasturbine engine.

The turbomachine aerofoil may be a turbine blade, a turbine vane, acompressor blade or a compressor vane.

The turbomachine aerofoil may be cast from a gamma titanium aluminide.

The curvatures of the convex surface and the concave surface arearranged to produce a thickness distribution of the aerofoil and acurvature of the concave surface which locates substantially all thesurface irregularities in that region of the concave surface where inoperation aerodynamic separation, a slow flow or a stagnant flow occurs.

The present invention also provides a method of producing a turbomachineaerofoil comprising:

(a) producing a mould for casting the turbomachine aerofoil, the mouldhaving surfaces defining the convex surface and the concave surface ofthe turbomachine aerofoil,

(b) pouring molten metal, or molten alloy, into the mould,

(c) cooling the molten metal, or molten alloy, while within the mould toproduce a cast turbomachine aerofoil,

(d) placing the cast turbomachine aerofoil in an autoclave,

(e) applying heat and applying pressure to the cast turbomachineaerofoil while within the autoclave to remove any voidage within thecast turbomachine aerofoil and to produce surface irregularities on thecast turbomachine aerofoil,

wherein the producing of the mould includes arranging the mould surfacesto produce a thickness distribution of the turbomachine aerofoil and acurvature of the concave surface of the turbomachine aerofoil whichlocate any surface irregularities produced by the subsequent applicationof heat and pressure on the voidage within the cast turbomachineaerofoil only in that region of the concave surface of the turbomachineaerofoil where in operation aerodynamic separation, a slow flow or astagnant flow occurs.

Preferably the turbomachine aerofoil has a shroud, arranging the mouldto have a surface to define a recess in a surface of the shroud facingaway from the aerofoil.

Preferably the cast turbomachine aerofoil is removed from the mouldafter step (c) and before step (d).

The alloy may be a gamma titanium aluminide.

Preferably the turbomachine aerofoil is etched or blasted to remove anyreaction products formed between the metal, or alloy, and the mould.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will be more fully described by way of examplewith reference to the accompanying drawings, in which:

FIG. 1 is a partially cut away view of a gas turbine engine having aturbomachine aerofoil according to the present invention.

FIG. 2 is an enlarged longitudinal cross-sectional view through the gasturbine engine in FIG. 1 showing a turbomachine aerofoil.

FIG. 3 is a longitudinal cross-sectional view through a mould forproducing a turbomachine aerofoil according to the present invention.

FIG. 4 is a cross-sectional view in the direction of arrows A--A in FIG.3.

FIG. 5 is a cross-sectional view in the direction of arrows B--B in FIG.3.

FIG. 6 is a cross-sectional view through a cast turbomachine aerofoil.

FIG. 7 is a cross-sectional view through a cast turbomachine aerofoilafter hot isostatic pressing, and

FIG. 8 is a view in the direction of arrow C in FIG. 6 of the concavepressure surface of a cast turbomachine aerofoil after hot isostaticpressing.

DETAILED DESCRIPTION OF THE INVENTION

A gas turbine engine 10, shown in FIG. 1; comprises in flow series anintake 12, a fan section 14, a compressor section 16, a combustionsection 18, a turbine section 20 and an exhaust 22. The turbine section20 comprises a plurality of turbine rotors 24 arranged to drive the fan(not shown) of the fan section 14 and compressor rotors (not shown) ofthe compressor section 16 via shafts (not shown). Only one of theturbine rotors 24 of the turbine section 14 is shown. Each of theturbine rotors 24 carries a plurality of turbine blades 26. The turbinesection also comprises a turbine stator casing 28 which supports aplurality of stages of turbine vanes 30. Each stage of turbine vanescomprises a plurality of turbine vanes 30.

A turbine rotor 24 and its associated turbine blades 26 are shown moreclearly in FIG. 2. Each turbine blade 26 comprises a root portion 32, ashank portion 34, a platform portion 36, an aerofoil portion 38 and ashroud portion 40. The platform portion 36 and the shroud portion 40partially define the inner and outer boundaries of the flow path of thehot gases through the turbine section 20. The shroud portion 40 also hassealing fins 42 which cooperate with the turbine casing 28 to preventthe leakage of hot gases around the shroud 40. The root portion 32 isshaped to locate in correspondingly shaped grooves on the rim of theturbine rotor 24 in order to retain the turbine blade 26 on the turbinerotor 24. The aerofoil portion 38 has a concave pressure surface 44 anda convex suction surface 46, as shown more clearly in FIGS. 6 and 7. Theconcave pressure surface 44 and the convex suction surface 46 extendfrom the leading edge 52 to the trailing edge 54 of the aerofoil portion38.

There are new alloys suitable for use in the turbine blade 26 andturbine vanes 30 of the lower pressure and lower temperature stages ofthe turbine section 20. These new alloys, or intermetallics, are knownas titanium aluminides, particularly gamma titanium aluminides forexample Ti₄₈, Al₄₈, Mn₂, Nb₂ where the suffix denotes the atomic percentof the elements. A problem exists when these gamma titanium aluminideintermetallics are cast in moulds, because there is a relatively largeamount of voidage, or porosity, in the cast turbine blade, or castturbine vane, when compared to similar superalloy castings. The voidageis removed by hot isostatically pressing the gamma titanium aluminideintermetallic castings in an autoclave, however this results in surfaceirregularities, ie dimples, on the surface of the cast and hotisostatically pressed gamma titanium aluminide turbine blade, or turbinevane. These surface irregularities are normally removed by casting theturbine blades, or turbine vanes, oversize and then machining to sizeand shape, or by filling the surface irregularities by weld deposition.

The present invention is based upon the concept of arranging for anysurface irregularities 48, in the cast and hot isostatically pressedturbine blade 26, or turbine vane 30, to be solely in the region 50 ofthe concave pressure surface 44 where in operation aerodynamicseparation occurs, ie where there is a reverse flow or a recirculatingflow as is shown in FIG. 7, or where in operation there is a very slowflow or a stagnant flow next to the concave surface which is of similaror greater dimension to the depth of the surface irregularities andlarger than the boundary layer on the concave surface. This will havethe effect of minimising any aerodynamic effects that the surfaceirregularities 48 have on the gas flows around the turbine blade 26, orturbine vane 30. Thus it is not necessary to machine off or fill in thesurface irregularities 48. The gas flow D splits at the leading edge 52of the aerofoil portion 38 of the turbine blade 26 to flow E over theconvex suction surface 46 and to flow F over the concave pressuresurface 44. There is a separation region G on the concave pressuresurface 44 but the main gas flow F is reaccelerated and reattaches tothe concave pressure surface 44 with low aerodynamic loss.

In order to achieve this objective the curvatures of the convex suctionsurface 46 and the concave pressure surface 44 of the aerofoil portion38 of the turbine blade 26, or turbine vane 30, are arranged to producea thickness distribution of the aerofoil portion 38 and a curvature ofthe concave pressure surface 44 which locate the surface irregularities48 only in that region 50 of the concave pressure surface 44 where inoperation aerodynamic separation, a slow flow or a stagnant flow occurs.Any aerodynamic effects on the concave pressure surface 44 are dependentupon its curvature. Voidage, or porosity, effects are dependent upon thethickness distribution of the aerofoil portion 38. The thicknessdistribution of the aerofoil portion 38 depends upon the curvatures ofthe convex suction surface 46 and the concave pressure surface 44. Thusby adjusting the curvatures of the concave pressure surface 44 and theconvex pressure surface 46 substantially all the voidage in the castturbine blade will be in the region 50 where aerodynamic separation, aslow flow or a stagnant flow occurs.

The method of producing a turbine blade, or turbine vane, is shown inFIGS. 3 to 8. Initially a wax pattern 60 is produced to define thepredetermined size and shape of the turbine blade, or turbine vane. Thewax pattern 60 has a concave surface 62 and a convex surface 64corresponding to that required for the aerofoil portion 38 of thefinished turbine blade 26, or turbine vane 30. A ceramic shell mould 66is formed around the wax pattern 60 by immersing the wax pattern 60 in aliquid ceramic slurry which quickly gels after draining and bysprinkling strengthening refractory granules over the ceramic slurrycovered wax pattern 60 to produce a ceramic layer on the wax pattern 60.The process is repeated several times to produce a ceramic layer whichhas a total thickness of about 1/4 inch (6 mm) to 1/2 inch (12 mm) onthe wax pattern 60 as shown in FIGS. 3 to 5. The wax is then melted outleaving a ceramic shell mould 66 having an internal cavity identical inshape to that of the wax pattern 60. The ceramic shell mould 66 is firedat a high temperature between 950° C. and 1100° C. to purify it byremoving all traces of residual wax, while at the same time curing theceramic shell mould 66. The ceramic shell mould 66 is then transferredto a casting furnace. A charge of molten gamma titanium aluminide isthen poured into the ceramic shell mould 66 and the ceramic shell mould66 is allowed to cool to room temperature, after which the ceramic shellmould 66 is removed leaving the cast turbine blade 26, or turbine vane30, as shown in FIG. 6. The voidage, or porosity, 56 in the turbineblade 26 is only in the region 50, with respect to its chordal length,where aerodynamic separation, a slow flow or a stagnant flow will occur.

The cast turbine blade 26 is then placed in an autoclave, and theautoclave is sealed. The autoclave is evacuated and then inert gas, forexample argon, is introduced into the autoclave. The pressure applied onthe cast turbine blade 26 by the inert gas is increased andsimultaneously the temperature in the autoclave is increased. Thetemperature is increased to between 1150° C. and 1320° C. and thepressure is increased to between 69 MPa and 310 MPa. These temperaturesand pressures are then maintained substantially constant for severalhours, for example 4 hours, and then the temperature and pressure arereduced to ambient. The consolidated turbine blade 26 as shown in FIGS.7 and 8 has had the voidage, or porosity, removed, but surfaceirregularities 48 appear in the region 50 of the concave pressuresurface 44 where aerodynamic separation, slow flow or a stagnant flowoccurs in operation, the remaining portion of the concave pressuresurface 44 and the whole of the convex suction surface 46 aresubstantially free of surface irregularities.

In this example the turbine blade 26 has a shroud portion 40. The shroudportion has a recess 41 in its surface facing away from the aerofoilportion 38. The recess 41 is produced by defining a recess 65 in thecorresponding surface of the shroud of the wax pattern 60. The effectsof this recess 41 is to cause any voidage, or porosity, in the region ofthe aerofoil portion 38 nearest the shroud portion 40 of the castturbine blade 26 to move towards the recess 41 and to increase the sizeof the recess 41 during the hot isostatic pressing process.

After the hot isostatic pressing step the surface of the aerofoilportion 38 is etched, or blasted, to remove any reaction products formedbetween the gamma titanium aluminide and the ceramic shell mould. Theroot portion 32 and the shank portion 34 are machined all over.

Although the invention has been described with reference to a gammatitanium aluminide comprising 48 atomic % titanium, 48 atomic %aluminum, 2 atomic % niobium and 2 atomic % manganese, it is possible touse other gamma titanium aluminides comprising 40 to 52 atomic %titanium, 44 to 52 atomic % aluminium and one or more of chromium,carbon, gallium, molybdenum, manganese, niobium, nickel, silicon,tantalum, vanadium and tungsten in an amount of between 0.05 to 8 atomic%. These gamma titanium aluminides may have a titanium diboridedispersoid in an amount between 0.5 to 20% volume. The invention is alsoapplicable to any other metals, alloys or intermetallics which haverelatively large amounts of voidage, or porosity, when cast, and whichproduce surface irregularities when consolidated by hot isostaticpressing. The invention is also suitable for use on gas turbine enginecompressor blades or compressor vanes and also steam turbine blades. Theinvention is also applicable to turbomachine aerofoils which areunshrouded.

The invention enables the turbomachine aerofoil to be cast to exact sizeand shape without the need for weld filling of any surfaceirregularities, or dispenses with the need to cast oversize and tomachine off excess metal, or alloy, including the surface irregularitiesto the exact size and shape required.

I claim:
 1. A method of producing a turbomachine aerofoil having aconvex surface and a concave surface comprising:(a) producing a mouldfor casting the turbomachine aerofoil, the mould having surfacesdefining the convex surface and the concave surface of the turbomachineaerofoil, (b) pouring molten metal, or molten alloy, into the mould, (c)cooling the molten metal, or molten alloy, while within the mould toproduce a cast turbomachine aerofoil, (d) placing the cast turbomachineaerofoil in an autoclave, (e) applying heat and applying pressure to thecast turbomachine aerofoil while within the autoclave to remove anyvoidage within the cast turbomachine aerofoil and to produce surfaceirregularities on the cast turbomachine aerofoil, wherein the producingof the mould includes arranging the mould surfaces to produce athickness distribution of the turbomachine aerofoil and a curvature ofthe concave surface of the turbomachine aerofoil which locate anysurface irregularities produced by the subsequent application of heatand pressure on the voidage within the cast turbomachine only in thatregion of the concave surface of the turbomachine aerofoil where inoperation aerodynamic separation, a reverse flow, a slow flow or astagnant flow occurs.
 2. A method as claimed in claim 1 in which theturbomachine aerofoil has a shroud, arranging the mould to have asurface to define a recess in a surface of the shroud facing away fromthe aerofoil.
 3. A method as claimed in claim 1 in which the castturbomachine aerofoil is removed from the mould after step (c) andbefore step (d).
 4. A method as claimed in claim 1 in which the alloy isa gamma titanium aluminide.
 5. A method as claimed in claim 1 in whichthe turbomachine aerofoil is etched or blasted to remove any reactionproducts formed between the metal, or alloy, and the mould.